Embodiments of the present disclosure relate to anti-icing systems for aircraft jet engine (i.e. gas turbine) propulsion systems and more particularly, to nozzle configurations on an injector in a swirling rotational anti-icing system for an inlet of a gas turbine engine.
Safety is a primary concern in the design of power propulsion systems for aircraft applications. The formation of ice on aircraft wings, tail fins, flight control surfaces, propellers, air inlets of engines, etc., has been a problem since the earliest days of heavier-than-air flight. Any accumulated ice adds considerable weight, and changes the airfoil or inlet configuration thereby making the aircraft much more difficult to fly and, in some cases, has caused loss of aircraft. In the case of gas turbine aircraft, large pieces of ice breaking loose from the leading edge of the gas turbine engine inlet housing can severely damage rotating fan-, compressor-, and turbine blades which extend across the flow path and interact with the working fluid or other internal fixed engine components (i.e. stator vanes) and cause engine failure.
Many attempts have been made to overcome the problems and dangers of aircraft icing. Anti-ice systems for the inlet area of nacelles for aircraft propulsion systems have been the focus of a significant amount of research and development within the aircraft industry. For example, proposals have been made, as described in U.S. Pat. No. 2,135,119 to mechanically vibrate external surfaces to break ice loose or as described in U.S. Pat. No. 3,549,964 to generate electromagnetic pulses in the aircraft skin to break ice loose. These systems, however tend to be heavy and complex and remove only existing ice, rather than prevent ice formation.
Heating areas of the aircraft prone to icing has been suggested many times. The heating schemes suggested range from microwave heating as suggested by U.S. Pat. No. 4,060,212 to feeding hot gases through holes in the skin, as suggested by U.S. Pat. No. 4,406,431, to resistance heating of the surfaces, as in U.S. Pat. No. 1,819,497, to actually burning fuel adjacent to ice-prone surfaces, as described in U.S. Pat. No. 2,680,345. While each of these methods may have some advantages none had been truly effective.
One of the most common anti-icing techniques has been the ducting of hot gases into a housing adjacent to the likely icing area. Typical of the patents describing such hot gas techniques are U.S. Pat. Nos. 3,057,154; 3,925,979; 3,933,327 and 4,250,250. In each case, the hot gas conduits simply dump hot gases into the housing, such as the leading edge of a gas turbine engine housing or a wing leading edge. While often useful, these systems are not fully effective due to the complexity of the hot gas duct system.
A typical design for a civil aircraft gas turbine engine nose cowl ice protection system was the double skin, spray bar configuration which employed an annular duct or skin installed within the nose cowl “D-duct” or D-shaped duct space. Hot bleed air jets issuing from small orifices, or piccolo holes, on the spray tube were directed on entrance into the D-duct into double skin passages along the upper and lower lip surfaces of the nose lip of the inlet. Hot bleed air was then forced to flow through narrow gaps between the outside skin and the inner skin, transferring heat to the outer skin. While some heat effectiveness was achieved by the double skin design it was at the expense of high manufacturing cost and weight penalty, associated with the required chem-milling process to produce the inner skin and to provide the spray tube arrangement.
Another advance in anti-icing systems was made in U.S. Pat. No. 4,688,745 entitled “Swirl Anti-Ice System” and issued to Rosenthal. This patent provided for the circulation of heated gases within the circular leading edge of a gas turbine engine housing in a rotational swirling motion to prevent ice build-up thereon. Hot gas, such as bleed air from a hot, high-pressure section of the gas turbine engine compressors was directed to the D-duct inlet interior through a conduit that enters the annular leading edge housing through a bulkhead closure. The conduit is then turned about 90° to a direction tangential to the leading edge annulus. The hot bleed air exiting an injection nozzle, provided as an outlet of such conduit entrain the cooler D-duct rotating air in the circular leading edge and cause the much larger mass of cooler D-duct rotating air within the inside of the inlet lip to swirl circularly around the interior of the D-duct of the annular housing. The entering hot bleed air heats the mass of cooler D-duct air to an intermediate but still relatively high temperature which then uniformly transfers heat to the skin of the leading edge without leaving any relatively cold areas and thereby preventing the formation of ice on the inlet lip. A fraction of the flow of entrained heated air that is equal to the flow rate of injected hot bleed air is exhausted from such D-duct housing.
While this anti-icing system represented a significant advancement over the prior art and has been widely accepted in the aircraft industry, there are areas of the system which may be improved. It has been found that as the near-sonic or sonic jet nozzle injects the hot bleed air into the cooler D-duct rotating air within the annular nose cowl or nose lip of the inlet, the complete mixing of the two masses of air is somewhat delayed during the rotational swirling action and results in a “hot spot” or area of elevated temperature on the outer lip skin of the nose lip at a position downstream of such injection. This area of elevated temperature in the nose lip then presents a constraint in the design of an anti-icing system according to teachings of such patent since such design must take into account actual conditions such as a day having an elevated ambient temperature, a low altitude take-off location for the aircraft, a high engine power setting, and a failure to an open position of an anti-ice valve, provided in the conduit for supplying hot bleed air from the engine's compressors as required.
Thus, there is a continuing need to improve aircraft engine nacelle icing prevention and to improve particularly the anti-icing system of U.S. Pat. No. 4,688,745 by enhancing the mixing of the injected hot gas and the mass of swirling air contained within the D-duct nose lip of an aircraft jet engine housing to improve the performance of the anti-icing system and to lessen design constraints imposed by the area of elevated temperature in the nose lip downstream of the point of injection of the hot bleed air.